A number of propulsion systems have been studied for use in hypersonic cruise and transatmospheric vehicles. These include, for example, the dual mode ramjet/scramjet engine and the rocket-based combined cycle (RBCC) engine. Unfortunately, each conventional propulsion system incorporates limitations that preclude the adoption of any single propulsion system suitable for all applications.
A dual mode ramjet-scramjet engine, for example, has good specific impulse over a Mach number range from about 3 to 10, but does not operate efficiently at either lower or higher Mach numbers, and has no static thrust capability.
The use of Rocket Based Combined Cycle (RBCC) propulsion systems could potentially offer improved performance over that of conventional cycles involving purely rocket or purely air-breathing concepts. Daines and Segal provide a broad summary of recent research in this area, with emphasis on weight savings, flight path planning and cycle efficiencies. The individual rocket and air-breathing cycles have inherent limitations. Although capable of operating over a very wide range of altitudes and Mach numbers, the pure rocket mode is hampered by very low specific impulse compared to air-breathing engine concepts. Air-breathing engines yield substantially higher values of specific impulse but conventional engine concepts are only capable of efficient operation over a narrow Mach number and altitude range. By combining the attributes of both rocket and air-breathing engine concepts, the RBCC offers improvements in performance.
Unfortunately, operation of RBCC engine concepts are inherently limited by the energy release rates of the deflagrative combustion processes employed in their combustion chambers. Deflagration combustion is a type of burning in which the flame front propagates into the fuel-oxidizer mixture at low subsonic speeds. This imposes limitations on energy release rate, and is accompanied by losses in total pressure.
More recently, pulse detonation engines (PDEs), operating in either an air-breathing or rocket mode, have been proposed as an alternative to conventional propulsion systems. The pulse detonation engine is an intermittent, constant-volume combustion engine concept. Air breathing engines based on the PDE cycle offer the promise of improved cycle efficiency and specific thrust, reduced specific fuel consumption and a wide operational range, while improved specific impulse has been demonstrated for certain rocket engine applications. Both propulsion systems may benefit from significant weight and cost reductions due to reduction or elimination of turbomachinery components.
A pulsed detonation engine consists of a chamber in which fuel and oxidizer are mixed and filled followed by a spark ignition. This initiates a detonation wave, which is a complex wave system generally characterized by a sharp pressure rise (up to 30 times or more), and a region of chemical reaction leading an expansion wave. This wave system traverses the detonation tube and is finally allowed to expand into a nozzle system. The high-speed combustion approximates a constant volume process and results in greater cycle efficiency. The detonation tube thus accommodates compression, combustion and expansion in one simple flow path.
Unfortunately, the current concepts being advocated for pulse detonation engines have performance limitations when required to operate over a wide Mach number range. The need to reduce the Mach number from its free stream value to a very low value prior to entering the detonation chamber through the intermittent air valve generally prevents the operation of air-breathing PDE's above flight Mach numbers of about 4. Above Mach 4, the temperature rise associated with the gas dynamic compression process in the inlet causes the air temperature to exceed the autoignition temperature of the fuel, which prevents detonation from occurring. Pulse Detonation Rockets (PDR) avoid this problem, but their specific impulse capability at low Mach numbers is much lower than air-breathing PDE's.